1. Field of the Invention
The present invention relates, generally, to gas turbine engines. More particularly, the invention relates to compressors for gas turbine engines for aircraft. The invention has particular utility for improving the efficiency of the high-pressure compressor on aircraft gas turbine engines.
2. Description of the Prior Art
Though it does not depict any existing engine, FIG. 1 illustrates the current state-of-the-art for gas turbine high-pressure compressor spools 10 in aircraft engines, and is included to provide a frame of reference for the subsequent discussion of prior art and for the present invention.
Efforts are ongoing to reduce fuel burn per pound of thrust in gas turbine engines for aircraft. One way to achieve that is to increase the efficiency of the compressor. There are several approaches for achieving that: (1) increase the maximum surface speed of the spool; (2) improve the efficiency of each stage of the compressor; and (3) increase the compressor pressure ratio capability.
Material stress limits and stiffness of conventional spool structures do not allow for increasing the maximum surface speed without adding significant mass to the spool structures. Current spool speed limits are set by rapid radial excursions from idle to maximum power conditions at the spacer mid-span locations between adjacent disks. For example, those conditions would occur if a runway incursion forced an aircraft to take off again immediately after touching down. At those conditions, the stator is at near idle temperatures and the spool is near its maximum speed and temperatures, resulting in large radial deflections and life limiting low cycle fatigue (LCF) stresses at the mid-span of the disk spacers 12, 14, and 16, which typically is where the seal teeth, 18, 20 and 22 are located. The seal teeth exacerbate the LCF stress limit problem because of local concentration of stresses at the teeth, and because of hard abrasive coatings that are frequently used to coat the tips of the teeth.
To minimize the weight of the compressor spool structure, rotor disks 24, 26, 28, and 30 are integral with blades 32, 34, 36, and 38 respectively for compressors used in aircraft engines. Disks and spacers are designed to minimize the number of bolts 40, 42 required to assemble the spool 10. For example, disk 26 has spacers 12 and 14 integrally attached, such as by welding, and disk 30 has spacer 16 and rear stub shaft 44 integrally attached. Such a practice makes disks 26 and 30 with all their integral structure very complex and expensive to make, repair, or replace.
Rotor disks were not always bolted together as illustrated in FIG. 1. In the 1950s, the J93 engine made by General Electric used a tie rod arrangement to hold the disks together. At the time, the airfoils used in the high-pressure compressor (HPC) for the J93 had a much higher aspect ratio than those of contemporary HPC's, and the resulting limited axial space between disks dictated a tie rod arrangement with spool external access to the threaded tie rods at assembly. The J93 engine also used curvics at the tie rods to centralize the disks and to provide a spool axial preload and torque path. Curvics, with their mating teeth, are expensive and heavy and they interfere with flow of secondary air within the spool. As the aspect ratio of airfoils used in compressors has decreased, the spacing between disks has increased, thereby allowing access for bolts between disks, and the configuration illustrated in FIG. 1.
The efficiency of each stage of the compressor can be improved by minimizing the cavity volumes between the stator shrouds 46, 48, and 50, and the disk spacers 12, 14, and 16. Circumferential air circulations in these cavities have an adverse impact on the compressor efficiency and stall margins. But conventional conical or cylindrical shaped spacers 12, 14, and 16 are not conducive to minimizing volume between them and substantially horizontal radial seal surfaces at the bottom of the shrouds 46, 48, and 50. An additional hindrance to minimizing volume between the first stage stator shroud 46 and the spacer 12 is the desire that the first stage stator vanes 52 be selectively pivotable about their axis. Since the first stage spacer 12 has a steep incline, significant clearance must be left for pivoting the stator vanes 52.
Compressor efficiency can also be improved by increasing the effectiveness of the compressor discharge pressure (CDP) seal 54 to minimize leakage. However, in conventional designs using a radial seal as shown, sufficient clearance, and hence, leakage must be allowed to accommodate radial excursions due to thermal cycling between static and maximum load conditions. Because the CDP seal 54 is located on the rear stub shaft 44, and the rear stub shaft 44 is attached at the rim of the last stage disk 30, the rear stub shaft 44 must be cooled.
Improvements in compressor performance will result in higher operating temperatures of the compressor, particularly at the last stage. Materials currently available limit the compressor discharge temperature to approximately 1150.degree. F., a value that has not increased by more than approximately 50.degree. F. for the last forty years. Because of this limitation, airfoil and disk life limits are reduced, particularly on the last stage of the compressor, when compressor performance improvements have been implemented. The airfoil and disk life limits can be increased by cooling the rear stages.
Several prior patents show apparatus and methods to cool the rotors of a gas turbine engine. U.S. Pat. No. 5,685,158 to Lenahan et al. discloses a means for extracting a portion of air from the combustor casing in a land-based gas turbine, routing it through a heat exchanger to cool it, and delivering the cooled air back to the rear of the compressor spool. U.S. Pat No. 5,897,386 to Kervistin discloses a means for tapping a portion of the bypass air for use in cooling the outer portion of rotor disks of the compressor and the base of blades attached to them. U.S. Pat. No. 5,755,556 to Hultgren et al. discloses a system of ducts in the rotors through which cooling air is directed. That system is designed for rotor disks with separately attached blades, as is the Kervistin system. U.S. Pat. No. 3,647,313 to Koff discloses a system for ducting air from the first compressor stage to the downstream end of the spool and back into the compressor inlet, with the air convectively circulating between rotor disks during its travel. This approach cools the aft disk bore portions and decreases their transient thermal response times. U.S. Pat. No. 4,793,772 to Zaehring and U.S. Pat. Nos. 4,920,741 and 4,961,309, both to Liebi, disclose circulating cooling air in a chamber formed outside of the stub shaft to cool the last compressor section. U.S. Pat. No. 4,808,073 to Zaehring et al. discloses vane-like ribs on the inside of the rear stub shaft which direct cooling air from the center shaft outwardly along the stub shaft and against the outer portion of the last rotor disk.
Another approach to disk transient thermal time reduction and cooling is disclosed in U.S. Pat. No. 4,648,181 to Putnam et al. Air is selectively bled from two different stages of the compressor, one warmer one cooler, and directed to the bore of the spool to actively control the rate of heating or cooling of the rotor, and thereby actively control and minimize the clearance of radial seals between the rotating and stationary portions of the compressor.
All of these active air methods require a source of air that has a high enough pressure to flow properly through the system, and a low enough temperature to provide the desired thermal effects. In aircraft engines, air taken from the bypass stream is very cool since it is not highly compressed, but its pressure may be insufficient to ensure proper flow direction, especially at low-power settings. Air taken from a high compression stage of the compressor has sufficient pressure, but may be too hot to provide effective thermal conditioning.
Recent advances in heat exchanger technology for aircraft gas turbine engines, specifically the use of phase change heat exchangers, have reduced the size of heat exchangers to practical sizes that allow high-pressure air from a high compression stage of the compressor to be sufficiently cooled to be effective in thermally conditioning the compressor spool components. The present invention makes use of such a source of thermal conditioning air and provides an improved high-pressure compressor spool which overcomes the limitations and shortcomings of the prior art.